1. Field of the Invention
The present invention concerns method and apparatus for in-flight acoustical monitoring of structural components in aircraft and more particularly is directed to a method and apparatus for measuring the actual remaining useful life of interchangable dynamic structural aircraft components, particularly helicopter rotor blades.
2. State of the Prior Art
It is known in the art to monitor acoustic emissions emitted by structures subject to loads in order to detect cracking or other failure in the structure. Typical instrumentation for carrying out such acoustic emission monitoring is disclosed by the following patents, which are primarily concerned with the transducer elements necessary to convert mechanical vibration to an electrical output for processing by suitable electronic circuits:
______________________________________ U.S. Pat. No. Patentee Issued Date ______________________________________ 3,774,443 Green et al November 27, 1973 3,855,847 Leschek December 24, 1974 3,779,071 Thomas Jr. et al December 18, 1973 3,529,465 Kleesattel et al September 22, 1970 4,088,907 Jones et al May 9, 1978 ______________________________________
A typical electronic circuit for processing acoustic emission signals is disclosed in U.S. Pat. No. 3,924,456 issued to Vahaviolos Dec. 9, 1975.
The need to monitor aircraft structures in order to detect or anticipate structural damage or failure before the aircraft ceases to be flightworthy has long been recognized as exemplified by U.S. Pat. No. 3,387,120 to Funk et al issued June 4, 1968 and U.S. Pat. No. 3,596,269 to Laska issued July 27, 1971. Both these patents disclose electronic or electro mechanical monitoring systems. Laska uses resistive components which are applied to critical surfaces such that the resistive components break upon failure of the structural member of interest. The resulting change in resistance is detected by appropriate circuitry and an indication of such failure is made to the aircraft crew. Funk et al discloses a system which measures acceleration loads imposed on the air frame and alerts the crew to excessive stresses. A circuit similar in principle to that of Laska is also disclosed in U.S. Pat. No. 4,106,332 issued to McKeown on Aug. 15, 1978.
The use of acoustic monitoring for testing of aircraft structures is disclosed as early as May 19, 1936 in U.S. Pat. No. 2,040,874 issued to Pack. The system disclosed therein, however, is not suited for continuous in-flight monitoring of the structure and is further limited to measuring resonant frequencies of structural members in response to an artificial mechanical stimulus which excites the member into a vibration mode.
The need to monitor the structural integrity and fatigue damage in helicopter rotor blades has been also recognized and solutions proposed in U.S. Pat. No. 3,744,300 issued to Fleury on July 10, 1973 and also in U.S. Pat. No. 3,985,318 issued to Dominey et al on Oct. 12, 1976. Fleury uses a resistive sensor for measuring progressive fatigue damage experienced by the rotor blade while Dominey et al uses a pressure differential sensing approach for use with hollow rotor blades the interior of which is at a pressure other than atmospheric such that cracks may be detected by the fact that the interior of the blade will be at atmospheric pressure. These two approaches suffer from obvious deficiencies in that Fleury is limited to sensing damage in the immediate area to which the resistive sensor is applied and would ignore changes in the structure occurring in areas removed from the sensor. Dominey et al is limited to blades in which a pressure differential can be maintained and is further deficient in that only an actual crack through the skin of the rotor blades sufficient to cause pressure equalization is capable of being detected. Progressive damage short of such a crack is not susceptible to detection by this system. The Fleury system is based on a determination of the number of force cycles to which the blade is subjected and on the amplitude of the force experienced in each cycle. This type of approach cannot detect progressive failure of a structural member occurring, for example, as a result of moisture infiltrating the interior of the structural member and causing degradation of adhesive bonds or corrosion of metallic structures. This kind of weakening is not a result of forces to which the blade is subjected and, therefore, would escape detection with this type of system.
A continuing need, therefore, exists for an in-flight monitoring system capable of monitoring progressive fatigue damage and deterioration of dynamic aircraft components and predicting actual failure of the structure prior to the occurrence of gross faults such as cracks required to trigger the Dominey et al system.
The need for reliable in-flight monitoring systems has become particularly acute in recent times with increasing use of composite materials in aircraft structures which are assembled through methods not yet fully tested, as opposed to welding, riveting and time proven assembly methods. Typically, composite materials are assembled by high temperature curing of epoxies which involve processes relatively novel to the aircraft industry and thus requiring the highest possible level of care to avoid accidents caused by unpredictable failure of structures for which no reliable life expectancy base lines have been established.